NACA 63A210 (naca63a210-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 63A210 (naca63a210-il) Reynolds number: 200,000 Max Cl/Cd: 48.26 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca63a210-il-200000-n5.txt Download as CSV file: xf-naca63a210-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A210
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6245 0.08895 0.08388 -0.0145 0.5422 0.0144
-10.000 -0.6338 0.08065 0.07566 -0.0201 0.5421 0.0143
-9.750 -0.6483 0.07005 0.06502 -0.0297 0.5421 0.0141
-9.500 -0.6658 0.06248 0.05733 -0.0367 0.5421 0.0139
-9.250 -0.6827 0.05700 0.05165 -0.0393 0.5421 0.0139
-9.000 -0.6935 0.05111 0.04542 -0.0405 0.5421 0.0140
-8.750 -0.6963 0.04584 0.03964 -0.0407 0.5420 0.0145
-8.500 -0.6951 0.04115 0.03454 -0.0405 0.5419 0.0151
-8.250 -0.6825 0.03869 0.03184 -0.0402 0.5418 0.0156
-8.000 -0.6683 0.03589 0.02869 -0.0397 0.5417 0.0160
-7.750 -0.6521 0.03313 0.02552 -0.0391 0.5416 0.0164
-7.500 -0.6334 0.03073 0.02274 -0.0385 0.5415 0.0169
-7.250 -0.6129 0.02862 0.02028 -0.0379 0.5413 0.0175
-7.000 -0.5909 0.02673 0.01806 -0.0373 0.5412 0.0182
-6.750 -0.5677 0.02521 0.01624 -0.0368 0.5411 0.0193
-6.500 -0.5432 0.02419 0.01496 -0.0363 0.5410 0.0207
-6.250 -0.5183 0.02318 0.01371 -0.0359 0.5407 0.0214
-6.000 -0.4955 0.02125 0.01160 -0.0353 0.5405 0.0224
-5.750 -0.4716 0.02015 0.01040 -0.0348 0.5402 0.0234
-5.500 -0.4471 0.01932 0.00950 -0.0344 0.5399 0.0247
-5.250 -0.4223 0.01863 0.00871 -0.0340 0.5397 0.0263
-5.000 -0.3970 0.01807 0.00807 -0.0337 0.5396 0.0289
-4.750 -0.3718 0.01748 0.00733 -0.0333 0.5395 0.0306
-4.500 -0.3476 0.01665 0.00643 -0.0329 0.5394 0.0334
-4.250 -0.3221 0.01611 0.00583 -0.0326 0.5394 0.0379
-4.000 -0.2960 0.01570 0.00532 -0.0323 0.5393 0.0441
-3.750 -0.2702 0.01522 0.00483 -0.0321 0.5392 0.0592
-3.500 -0.2450 0.01464 0.00442 -0.0319 0.5392 0.0992
-3.250 -0.2218 0.01370 0.00402 -0.0317 0.5391 0.2219
-3.000 -0.2003 0.01267 0.00375 -0.0312 0.5391 0.4094
-2.750 -0.1761 0.01225 0.00366 -0.0306 0.5391 0.5100
-2.500 -0.1508 0.01203 0.00358 -0.0300 0.5390 0.5713
-2.250 -0.1254 0.01189 0.00357 -0.0293 0.5390 0.6262
-2.000 -0.1001 0.01183 0.00361 -0.0286 0.5386 0.6726
-1.750 -0.0741 0.01184 0.00365 -0.0280 0.5380 0.7043
-1.500 -0.0473 0.01188 0.00365 -0.0277 0.5373 0.7250
-1.250 -0.0202 0.01192 0.00365 -0.0274 0.5367 0.7401
-1.000 0.0071 0.01195 0.00365 -0.0273 0.5363 0.7531
-0.750 0.0345 0.01201 0.00369 -0.0271 0.5357 0.7659
-0.500 0.0618 0.01204 0.00371 -0.0270 0.5350 0.7790
0.000 0.1161 0.01205 0.00376 -0.0266 0.5333 0.8060
0.250 0.1432 0.01207 0.00382 -0.0263 0.5325 0.8200
0.500 0.1705 0.01210 0.00390 -0.0261 0.5318 0.8351
0.750 0.1981 0.01215 0.00401 -0.0259 0.5310 0.8519
1.000 0.2265 0.01230 0.00420 -0.0260 0.5297 0.8704
1.250 0.2561 0.01264 0.00457 -0.0263 0.5279 0.8910
1.500 0.2882 0.01280 0.00479 -0.0272 0.5267 0.9123
1.750 0.3227 0.01290 0.00496 -0.0286 0.5253 0.9343
2.000 0.3577 0.01324 0.00536 -0.0303 0.5225 0.9556
2.250 0.3927 0.01396 0.00615 -0.0322 0.5165 0.9749
2.500 0.4274 0.01477 0.00704 -0.0341 0.5088 0.9923
2.750 0.4571 0.01550 0.00788 -0.0350 0.5013 1.0000
3.000 0.4813 0.01506 0.00768 -0.0344 0.4992 1.0000
3.250 0.5086 0.01210 0.00493 -0.0333 0.4611 1.0000
3.500 0.5337 0.01233 0.00496 -0.0330 0.4430 1.0000
3.750 0.5585 0.01267 0.00535 -0.0326 0.4266 1.0000
4.000 0.5846 0.01263 0.00556 -0.0322 0.4040 1.0000
4.250 0.6091 0.01262 0.00540 -0.0316 0.3237 1.0000
4.500 0.6290 0.01324 0.00544 -0.0306 0.2012 1.0000
4.750 0.6479 0.01423 0.00586 -0.0296 0.1001 1.0000
5.000 0.6698 0.01492 0.00634 -0.0288 0.0615 1.0000
5.250 0.6928 0.01549 0.00690 -0.0281 0.0461 1.0000
5.500 0.7157 0.01609 0.00752 -0.0274 0.0378 1.0000
5.750 0.7381 0.01673 0.00818 -0.0266 0.0318 1.0000
6.000 0.7598 0.01744 0.00896 -0.0257 0.0280 1.0000
6.250 0.7819 0.01811 0.00972 -0.0248 0.0257 1.0000
6.500 0.8034 0.01886 0.01055 -0.0239 0.0240 1.0000
6.750 0.8242 0.01969 0.01148 -0.0229 0.0226 1.0000
7.000 0.8437 0.02070 0.01254 -0.0217 0.0215 1.0000
7.250 0.8609 0.02212 0.01403 -0.0202 0.0206 1.0000
7.500 0.8796 0.02354 0.01555 -0.0189 0.0200 1.0000
7.750 0.9001 0.02481 0.01696 -0.0179 0.0197 1.0000
8.000 0.9209 0.02613 0.01845 -0.0169 0.0191 1.0000
8.250 0.9416 0.02735 0.01989 -0.0160 0.0182 1.0000
8.500 0.9616 0.02868 0.02140 -0.0150 0.0172 1.0000
8.750 0.9803 0.03041 0.02336 -0.0139 0.0167 1.0000
9.000 0.9975 0.03230 0.02550 -0.0126 0.0162 1.0000
9.250 1.0129 0.03433 0.02779 -0.0113 0.0157 1.0000
9.500 1.0260 0.03652 0.03025 -0.0098 0.0154 1.0000
9.750 1.0369 0.03873 0.03272 -0.0082 0.0150 1.0000
10.000 1.0463 0.04074 0.03492 -0.0067 0.0145 1.0000
10.250 1.0501 0.04343 0.03781 -0.0049 0.0140 1.0000
10.500 1.0430 0.04724 0.04196 -0.0025 0.0136 1.0000
10.750 1.0358 0.04990 0.04499 0.0002 0.0133 1.0000
11.000 1.0230 0.05325 0.04863 0.0022 0.0132 1.0000
11.250 1.0067 0.05731 0.05297 0.0029 0.0130 1.0000
11.500 0.9884 0.06224 0.05813 0.0019 0.0130 1.0000
11.750 0.9687 0.06830 0.06437 -0.0014 0.0130 1.0000
12.000 0.9478 0.07563 0.07188 -0.0066 0.0131 1.0000
12.250 0.9255 0.08379 0.08019 -0.0122 0.0132 1.0000
12.500 0.9020 0.09329 0.08979 -0.0187 0.0134 1.0000
12.750 0.8776 0.10418 0.10075 -0.0256 0.0137 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 63A210 (naca63a210-il)